Universal automatic landing system for remote piloted vehicles

ABSTRACT

An automatic landing system for landing remotely piloted flying vehicles  ng a predetermined path and at a predetermined point. The system includes an autopilot carried by the flying vehicle for measuring the parameters of attitude, airspeed, and heading and for comparing the measured parameters with the inputted parameters for the desired attitude, airspeed and heading. The autopilot adjusts the vehicle controls to make it conform to the desired attitude, airspeed and heading when deviations therefrom are detected. The system includes a radar transmitter and receiver means disposed on a stabilized double gimbal for measuring the actual heading and distance from the vehicle to the radar transmitter and receiver on a continuous basis. Control means are provided for receiving signals from the radar transmitter and receiver, computing actual and desired angles and altitude and for comparing hte actual parameters with the desired parameters and for instructing the autopilot in overcoming any deviations detected therein.

DEDICATORY CLAUSE

The invention described herein may be manufactured, used, and licensedby or for the U.S. Government for governmental purposes without thepayment to me of any royalties thereon.

BACKGROUND OF THE INVENTION

This invention relates to an automatic landing system which can beincorporated into any remotely piloted vehicle system where the aircraftor vehicle can be landed on wheels or skids or the system can beutilized to guide aircraft into a recovery net.

This system requires no addition to the remotely piloted vehicle andonly minor interface connections with the ground control system for suchvehicles. The automatic landing system uses state-of-the-art hardwarewhich is available, and which is combined in a unique way to makelanding of the remotely piloted vehicles easier for unskilled operators,and is completely automatic.

In one known system for automatic landing or recovering of remotelypiloted aircraft or a remotely piloted vehicle, an electro-opticalsensor is provided on the landing net structure which may be either atelevision or a forward looking infrared device, or similar device whichtracks a beacon which is mounted on the nose of the aircraft. In thissystem the electro-optical sensor on the net structure tracks thebeacon, and any errors arising from an incorrect flight path, that is,the failure of the air vehicle to fly directly towards the net, are usedas commands, either manually or automatically, to the air vehiclethrough the ground control station to correct the aircraft's flightpath. One problem with this system is that each aircraft is required tohave a particular type of beacon and may become unreliable in weatherconditions of low visibility for the television or infrared sensors andcan be used only for recovering the remotely piloted vehicle in a net.

Another known type of instrument approach is that used with mannedaircraft (known as a ground control approach) that uses a precisionradar to track the aircraft's approach, that is, both its glideslope andits heading. This system uses a ground controller to interpret the radardisplay and to advise the pilot of the manned aircraft as to neededcorrections in the flight path. This system is not adaptable to remotelypiloted vehicles.

SUMMARY OF THE INVENTION

It is an object of the invention to provide an automatic landing systemfor remotely piloted flying vehicles.

It is a further object of the invention to provide an automatic landingsystem for landing remotely piloted flying vehicles which does notrequire modification of the remotely piloted vehicle.

It is yet another object of the invention to provide an automaticlanding system for remotely piloted flying vehicles which includes asensor for sensing the altitude and angular displacement of the remotelypiloted vehicle at any distance.

It is still another object of the invention to provide an automaticlanding system for remotely piloted flying vehicles which includes asensor device for sensing the altitude and angular displacement of theremotely flying vehicle along a path offset from said detector.

These and other objects, which will become apparent, are accomplished byan automatic landing system which comprises a narrow beam radarseeker/sensor unit which is mounted on a two-axis pitch and yaw axisgimbal which is used to track the remotely piloted air vehicle from apredetermined point in space from where the landing approach begins to apoint just above the earth where the unmanned aircraft touches down. Theradar seeker/sensor and its gimbal set measures the pitch and the yawangles of the tracking beam and the range to the air vehicle. Theseangles and the range are used to compute the height above the terrainand to determine whether the path of the aircraft conforms to thedesired path to bring the unmanned vehicle to the safe landing at thepredetermined touch down point.

The sensor/seeker and its associated microprocessor calculates errors inthe desired flight path which, in turn, are input to the automaticcontrols as instructions or corrections for the air vehicle's flightpath. The design characteristic of the air vehicle are used indetermining the predetermined glideslope for the space provided for thelanding of the aircraft. The necessary geometrical relationships whichthe seeker/sensor microprocessor solve to determine conformance to thedesired flight path, are derived from the measurements made by theseeker/sensor.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention will now be described in connection with the appendeddrawings, wherein:

FIG. 1 is a perspective view of the automatic landing system of theinvention;

FIG. 2 is a diagrammatic side view of the glideslope controls of theinvention;

FIG. 3 is a diagrammatic plan view of the yaw control for the automaticlanding system of the invention;

FIG. 4 is a perspective view of the radar seeker/sensor unit of theinvention;

FIG. 5 is a rear elevation view of the radar seeker/sensor unit shown inFIG. 4;

FIG. 6 is a side view of the radar seeker/sensor unit of FIG. 5 butshowing it mounted on a tripod rather than a platform;

FIG. 7 is a schematic diagram illustrating the operation of theinvention;

FIG. 8 is a logic flow diagram for the glideslope control system; and

FIG. 9 is a logic flow diagram for the heading control system.

DETAILED DESCRIPTION OF THE DRAWING

Referring now to FIG. 1 of the drawings, the automatic landing system ofthe invention comprises a remotely piloted air vehicle 10, a groundcontrol unit 12 and a stabilized radar seeker/sensor tracking unit 14.Each of the subsystems of the invention are bought and marketed, buttheir basic operation will be briefly described before describing theoperation of the automatic landing system of the invention.

The remotely piloted air vehicle 10 and the ground control unit 12comprise the major components of an unmanned aerial vehicle and may be aremotely piloted vehicle system such as that known as the PIONEER systembuilt by the Tadrian Corporation of Israel and AAI, Incorporated ofBaltimore, Md., the HERON 26, built by the Pacific Advanced EngineeringCompany of San Diego, Calif.; and the SKYEYE built by DevelopmentalScience Division of the G.E.C. Avionics, Inc. of Ontario, Calif. Othermanufacturers manufacture similar units but all of these operateessentially in the same manner. The air vehicle 10 is a miniature, fixedwing aircraft, usually powered by a propeller or a jet which contains anautomatic pilot (which will be described in more detail later) whichcomprises gyroscopes or other instruments to measure the altitude andheading of the aircraft and electronically responsive controls. Theautomatic pilot controls the aircraft to fly a desired altitude andheading.

There are numerous manifestations of these autopilots, all of whichoperate by the principal of measuring the parameters of interest, i.e.,attitude, airspeed, heading, altitude, and the like, and comparing themeasured parameters with the desired values for those parameters. Thedifference between the measured parameters and the desired parametersindicate an error which the autopilot uses to make a change in theaircraft performance to correct the error.

The remote operator, in the ground control unit 12, commands theaircraft by changing the reference parameters or by inputting a desiredaction. For example, the operator may command the aircraft to make aright turn and the auto pilot will determine what adjustments arerequired in the aileron, rudder, elevator, and the engine controls tocause the air vehicle to conform to the command. Commands from theground control unit 12, are transmitted by a radio frequency to areceiver in the unmanned aircraft.

The millimeter wave radar seeker/sensor tracking unit 14 is a miniatureradar transmitter/receiver which is capable of tracking an object (inthis case the flying air vehicle) and measuring both its angulardeviation from a reference axis and its range from thetransmitter/receiver. The technology for such devices is well known andthere are many companies and individuals engaged in designing andbuilding radar frequency sensor systems. These devices are commonly usedas missile seekers for accurately tracking targets, as sensors forremotely piloted vehicles, in manned aircraft in target acquisitiondevices, and for helicopter and other aircraft detection and tracking byair defense weapon systems.

The basic radar seeker/sensor 14 is mounted on a stabilized gimbal (asseen in FIGS. 4, 5, and 6) which allow it to track an object over a wideangular excursion, as will be described in detail hereinafter. Thesensor/seeker and its mount are disposed on a trailer, truck bed or aportable tripod so that it can be set up at the site of an intendedlanding.

Referring now to FIG. 5 it will be noted that the stabilized radarseeker/sensor and tracking unit 14 is mounted on a pedestal mount 16,which may be the flatbed of a truck or otherwise, on which is supporteda gimbal support shaft 18 by plurality a of machine bolts (not shown).Gimbal support shaft 18 has mounted on its end a yaw gimbal 20 and a yawtorque motor 22 which is adapted to pivot yaw gimbal 20 relative to thesupport shaft 18 in either a clockwise or a counterclockwise direction.

Mounted on the end of yaw gimbal 20 is a pitch gimbal 24 which comprisestwo upstanding arms disposed on either side of the radar seeker/sensordevice. Pitch gimbal 24 also comprises a pitch gimbal motor 26, whichpivots the radar seeker/sensor within the arms of supporting gimbal 24.The stabilized radar seeker/sensor tracking unit 14 homes in on theremotely piloted vehicle and generates a signal responsive to the angleof glideslope and the heading of the remotely piloted vehicle. Thesesignals are transmitted to a microprocessor and control mechanism 28,which may be supported by the pedestal 16, through a seeker cable 30.The microprocessor 28 generates signals and transmits them to thecontrol unit 12, by means of a control unit cable 32, where the signalsare utilized to adjust the flight path of the remotely piloted vehicleto make it conform to the flight path predetermined to be the optimumfor the landing of the remotely piloted vehicle under the circumstances,and in accordance with the vehicle's operational configuration.

As seen in FIG. 6, the gimbal supporting the radar seeker/sensor andtracking unit 14 may be mounted on a tripod 34 instead of the pedestalillustrated in FIGS. 4 and 5.

Referring now to FIGS. 2, 3, and 7, it will be noted that the remotelypiloted aircraft 10 is controlled by the ground operator through manualcontrol 38 and standard controls 36 through a data link transmitter 40which transmits instructions to the autopilot 44 through a data linkreceiver 42 until the remotely piloted vehicle reaches point A, as seenin FIG. 2. When the remotely piloted aircraft reaches point A, themanual controls of the aircraft may be overridden and the automaticsystem takes control of the aircraft and lands it. Alternately, theoperator may initiate the automatic system. As seen in FIGS. 2 and 3,the radar seeker/sensor and tracking unit is mounted on the gimbal apredetermined distance off the ground and spaced a predetermineddistance from the landing path of the remotely piloted vehicle. Theaircraft 10 is guided in an initial guideslope A-F until it reachespoint B. Upon reaching point B, the guideslope of the remotely pilotedvehicle is altered to a final glideslope B-D until it reaches a touchdown point C.

As seen in FIG. 3, the desired heading for the remotely piloted vehicle10 carries it through point A, B, and C which are parallel to a baseline and null axis, established by the control and microprocessor 28.

The operation of the invention will now be described in connection withFIGS. 7, 8, and 9, as well as FIGS. 2 and 3.

The air vehicle 10 is flown to the final approach waypoint or location Aby commands from the ground control unit or by pre-programmed commandsstored in the autopilot in the usual manner. In the vicinity of thisfinal approach fix (point A of FIG. 2), the air vehicle altitude,heading, and air-speed will all be as required to initiate landing forthe particular aircraft design. This is standard practice for remotelypiloted (or manned) aircraft and requires no special consideration forthe application of this invention.

Through geometrical considerations (to be discussed later) and desiredaltitude to initiate the final approach to landing, the position ofpoint A (final approach fix) with respect to the radar seeker/sensor isdetermined and the seeker is made to point at that position in space. Asthe air vehicle approaches point A, the seeker/sensor detects it and theremotely piloted vehicle 10 operator may initiate automatic tracking,usually by placing a "tracking gate" over the air vehicle return on theradar image.

From this event forward, all actions are automatic and independent ofthe human operator. Because the seeker/sensor can begin tracking the airvehicle prior to its reaching point A, it is able to measure parameters(angle φ and range r of FIG. 2) which allow determination of when pointA is reached (a specific distance from, and altitude above theseeker/sensor). When that determination is made, the command is given tothe air vehicle to begin its initial approach. Depending on theindividual remotely piloted vehicle design, this command might be"descend at 500 feet/minute on a heading of 30 degrees" or it might be"decrease power to 60 percent, increase pitch up angle by 10 degrees,maintain constant heading". Of course, many possibilities exist forvariations on this scenario and any command that will cause the aircraftto descend at the predetermined constant rate and heading will besatisfactory.

Stored within the landing system computer 28 memory are equations (to bediscussed later) and parameters for computing the desired altitude H forany range r from seeker/sensor to the air vehicle. Therefore, the seekermeasures actual range r and angle φ as the aircraft descends, theassociated computer calculates actual altitude h and compares it to thedesired value (as will be described in detail later). If the two valuesdo not agree, the command is given to the air vehicle (via the groundcontrol unit and data link) to increase or decrease the rate of descent.Therefore, the air vehicle is constrained to fly small perturbationsabout the desired glideslope so that it does not get low enough tocollide with obstacles but is brought to a position to land on therunway.

As the air vehicle nears the ground, its rate of descent must be slowedso that is will touch down with a relatively low forward velocity and alow downward velocity to prevent damage to the vehicle. Thus, at a pointin space (point B of FIG. 2) as determined by the air vehicle reaching apredetermined altitude (perhaps 30 feet or so), a preprogrammed commandwill be given to reduce the rate of descent by reducing the glideslopeangle. This glideslope for the final portion of the descent will be at amuch smaller angle, causing a slower rate of descent that the initialglideslope. The actual glideslope will again be a function of thecharacteristics of the particular air vehicle because the final velocityof the air vehicle is dependent upon its stall speed. These velocitiesare part of the normal operating characteristics of an air vehicle andare readily determined.

After the command is given (from the landing system via the groundcontrol unit and data link to the air vehicle) to change the rate ofdescent at point B (FIG. 2), the air vehicles adherence to the newglideslope is monitored by the seeker/sensor as before. When the airvehicle reaches a point just above the ground, (perhaps 3 feet or so) asindicated by point C in FIG. 2, the final command is given tosimultaneously reduce engine power to idle and raise the nose (increasethe pitch up angle) so the aircraft will touch down at the stall speedor just above, and then roll forward until its forward velocity isdissipated. The geometrical considerations, parameters to be measuredand relationships to be used in calculating the necessary glideslopefunctions will be discussed in more detail later.

The second function of the automatic landing system (in addition toglideslope control) is heading control. This is achieved in a similarmanner to the glideslope control except that the process is considerablysimplified. From FIG. 3, it is seen that the seeker/sensor unit is setup offset to one side of the desired landing path (or runway) by adistance "s" (which will typically be about 50 feet but could be more orless). In the case of heading control, it is simply desired to cause theair vehicle to fly a straight path from the initial approach point tothe point of intended landing or touch down. The seeker/sensor could beset up directly in the landing path but the implementation is morepractical if it is offset to one side or the other (as will be discussedlater). The seeker/sensor is initially set up and calibrated to have itsnull axis (in the horizontal direction) parallel to the desired approachpath (as shown in FIG. 3). Then, at any and all points in its approach,the seeker/sensor (which is tracking the air vehicle in both pitch andyaw) will measure angle α and the range r to the air vehicle. These canbe used to measure whether the air vehicle is on the flight path or tothe right or left of it. If an error or deviation exists, a command isgiven to make a small turn right or left (as required) to correct it.

The geometrical relationships that are employed in performing the aboveoperations will be discussed in the following paragraphs.

To understand the means by which the system measures and controls theglideslope or vertical descent of the air vehicle, refer to FIG. 2. Eachair vehicle will have a best approach angle glideslope which isdependent upon its individual design. This glideslope is characterizedby the angle θ which is the angle below horizontal (level) flight atwhich the air vehicle approaches the ground. Two right trianglesdescribe the desired glideslope relationships. The first is made up ofthe height h of the air vehicle above a reference plane (made byextending the runway or landing field altitude to a point beneath theair vehicle), the desired glideslope, and the distance from theintersection of the glideslope with the reference plane to a pointexactly below the air vehicle. This is shown on FIG. 2 as H_(a), D_(a),G with included angle θ. It is desired to maintain θ as a constant sothat for any distance D_(a), there is a proper value of H which iscalculated by:

    H.sub.a =tan θ D.sub.a                               (1)

Capital letters are used to denote the desired values (as opposed to themeasured values). Thus, as the aircraft approaches the runway, thedistance D_(a) decreases and it is desired for H_(a) to decreaseaccording to equation (1). There is a similar triangle (H_(b), D_(b), G'and containing θ') which describes the desired glideslope relationshipduring the near ground portion of the approach when the glideslope angleis reduced to lower landing shock to the aircraft. The relationship ofheight to distance from glideslope intersection with the ground is:

    H.sub.b =tan θ' D.sub.b                              (1a)

Note that this is the same form as equation 1 except it is for the finalglideslope triangle.

The second triangle describes the relationship between the radarsensor/seeker and the aircraft. This triangle is made up of thecenterline of the tracking beam r which is also the measured range fromthe seeker to the air vehicle; the height h of the air vehicle above thereference plane less the height of the radar seeker above the referenceplane (Δ); and the distance (d_(m)) from the radar seeker to a pointbeneath the air vehicle. The included angle phi(φ)is a variable that ismeasured from the gimbal angles of the seeker. These parameters areexpressed in the lower case to indicate measured parameters orparameters calculated from measurements. Thus, the measured height ofthe air vehicle (above the ground) is:

    h=sin φ r+Δ                                      (2)

Since height is only useful information when it is associated with adistance (from a known point, in this case the radar seeker/sensor), thedistance relationship is given by:

    d.sub.m =cos φ r                                       (3)

This distance (calculated from the measured data) is then related to thedistances in the glideslope triangles by:

    D.sub.a =d.sub.m -X                                        (4a)

    D.sub.b =d.sub.m -X'                                       (4b)

With the equations (and the known or arbitrary parameters), theinformation from the radar seeker/sensor (which is range r and angle φ)can be used to compute the height and distance (actual, as well asdesired height at the same distance). Among the arbitrary parameters areinitial altitude to begin the approach (H_(a) initial), glideslopeangle, altitude where final glideslope is initiated (H_(b) initial), andfinal glideslope angle. Although these are arbitrary, they are greatlyinfluenced by the design of the aircraft and the terrain (includingobstruction heights) surrounding the landing area. Also, the distance X'is arbitrarily chosen to allow for some overshoot in the approach, ifdesired.

Therefore, for any position of the air vehicle on or around theglideslope, a range r and an elevation angle φ are measured. Then usingequation 2 and 3, the height h at a distance d_(m) from theseeker/sensor is calculated. Equation 4 is used with the value of d_(m)and the arbitrary X or X' to determine a distance D_(a) or D_(b) fromwhich equation 1 is used to calculate the desired height (H_(a) orH_(b)). Then the desired height (H_(a) or H_(b)) is compared to theactual height h. If these two values are the same, no action is takenbecause the air vehicle is on the correct glideslope. If they areunequal, the following logic is implemented:

For h<H; add power and/or pitch up

h>H; reduce power and/or pitch down

Upon the air vehicle reaching point B (as indicated by its descent toheight H_(b) at distance D_(b)), the air vehicle will be commanded toreduce its glideslope angle to a new value (θ' in FIG. 2). The sameprocedure will be followed as during the initial glideslope phase tomeasure and correct for altitude (height) errors.

The logic to be used in making the transition at points A, B, and C isessentially the same. At any point in time (after the seeker/sensorbegins tracking the air vehicle), various computations are being madeand corrections made as required to conform the actual flight path tothe desired approach path. FIG. 8 is a flow diagram which illustrates,in general, how the logic and computation by the seeker/sensor computeris performed.

In order to better explain the process, a numerical example will now bedescribed. This example will refer to FIGS. 2 and 8. The first thing tobe done is to select those parameters that are arbitrary or related tothe performance of an individual air vehicle. The initial altitude(H_(ai)), which is the height of the aircraft at the beginning of theautomatic approach, is chosen to be 500 feet. The initial glideslopeangle θ is selected as 20 degrees. These are values that might typicallybe used but many other values could also be selected. With theseparameters chosen, the initial value of D_(a) (FIG. 4) is calculatedusing equation 1: ##EQU1##

Next the parameters of the final glideslope triangle (H_(b), D_(b), andθ' of FIG. 4) are considered. The initial value for H_(b) (H_(bi)) ischosen as 30 feet while the glideslope (θ') is chosen as 5 degrees. Thenthe initial value of D_(b) (D_(bi)) is calculated using equation 1(a):##EQU2##

From FIG. 2, it is observed that the initial and final glideslopetriangles overlap and the distance they overlap (D_(o)) is the distancefrom the intersection of the initial glideslope with the reference plane(F) and the point where the initial H_(b) value occurs (point B). Thisoverlap distance is found using equation 1, with a parametermodification (H_(a=H) _(bi) ; D_(a) =D_(o)). This is: ##EQU3##

The final arbitrary choice to be made is the distance from the projectedtouch down point D to the sensor/seeker E (distance X' on FIG. 2). Thiscould be zero if desired but it would normally be made 100-300 feet toallow for a slightly high approach (with longer than expected touchdown). Let X'=200 feet for this example. From FIG. 2 it is seen thatdistance X can be calculated from X', D_(bi), and D_(o) (all of whichhave been selected or calculated). Therefore:

    X=X'+D.sub.bi -D.sub.o

    X=200+343-82=461 feet

The only other parameters are the height of the sensor above thereference plane(Δ), which is chosen to be 3 feet and the height H_(ci)(which is the height where landing is imminent) and the aircraft enginepower is reduced to zero and the nose is raised to put the aircraft inthe final landing attitude. Tis will be chosen to equal Δ (3 feet).

Thus, all the baseline parameters have been selected and calculated. Itis obvious that these parameters can be varied to fit the existingconditions. For example, if there is a very small landing area, thevalue of H_(ai) can be made lower and the angle θ can be made steeper toallow landing in less space. On the other hand, a large, fast aircraftmay require a higher H_(ai) and a less steep angle to accommodate itsperformance limitations.

By referring to FIG. 8 as well as FIG. 2, the basic dynamic operation ofthe landing system can be described. On the upper right side of FIG. 8is shown the arbitrary or constant values that are selected prior to use(either by normally entering data from a handbook or by calling upappropriate data from a computer memory device). These parameters areH_(ai), H_(bi), θ, θ', Δ, and X', whose values were calculated above. Onthe upper left of FIG. 8 is shown the dynamic measured parameters φ andr from the radar seeker/sensor. Starting at the top of FIG. 8, the angleφ and r measurements are used to calculate (in C₁) (continuously) thedistance d_(m) (shown on FIG. 2) using equation 3. Meanwhile, theinitial value of distance D_(a) (that point where the automatic approachbegins) is calculated (in C₂) from the arbitrary values of H_(ai) andusing equation 1. Then, in the first decision block (D₁), the currentvalue of d_(m) is compared to D.sub. ai (which is calculated in C₂). Ifthese values are not equal, then the next decision block D₂ checks tosee if d_(m) is less than D_(ai) (d_(m) <D_(ai)). If the answer is no,the aircraft has not yet reached the point where the approach begins andno action is taken by the landing system. When d_(m) does equal D_(ai)the "yes" output of the first decision block D₁ sets the engine andpitch controls to their proper values for the initial glideslope. At thesame time switch S₁ is closed, which injects the H_(a) value(equation 1) into the logic for calculating and comparing height of theaircraft. Then as d_(m) becomes less than D_(ai) (which indicates thatthe aircraft has started down for the approach), the "yes" output of thesecond decision block D₂ is activated which causes the computation ofthe actual height of the aircraft to be calculated (C₃) (from equation2) using the dynamic parameters φ and r and the static parameter Δ. Theoutput of this computation (h) goes to several more decision blocks. Thefirst (D₃ on the diagram) tests to see if h is equal to or less thanH_(bi). During the phase of the approach between point A (where H_(ai)is encountered) and point B (the point where H_(bi) is encountered),this output will be "no" and no action will be taken as a result.

At this point it should be noted that on the right side of the logicdiagram, D_(o), D_(bi), and X are calculated (in C₄, C₅, and C₆) fromthe fixed (or arbitrary) parameters (as discussed previously) and usedalong with the glideslope angles (θ and θ') and d_(m) (calculated fromthe dynamic measurements of φ and r) to calculate the desired values ofH_(a) and H_(b) (heights above the reference plane in the two glideslopetriangles) in C₇ and C₈. Only one value of H (H_(a) or H_(b)) is used ata time and switches S₁ and S₂ determine which.

During the period when the aircraft is on the initial glideslope(between points A and B), a negative output of decision block D₃(h≦H_(bi)) causes no action and h is also equated with H_(a) is decisionblock D₄ (since switch S₁ is closed and S₂ is open). If H_(a) is equalto h (yes output of decision D₄), the aircraft is exactly on the desiredglideslope and no action is taken. If this equality is not true, (a "no"output of decision D₄) then decision D₅ checks to see if h is less thanH_(a). If this is true ("yes" output), then the aircraft is below theglideslope so the power is increased and pitch up angle is increased. Ifthis is not true ("no" output), the aircraft is above the glideslope sothe power is reduced and the nose is pitched down to lose altitude.

This process continues to maintain the aircraft position on theglideslope. Since d_(m) and h are being computed continuously,undesirable deviations in aircraft height are determined and correctedbefore they can become large.

When the aircraft height h equals H_(bi), the "yes" output of decisionD₃ (h≦H_(bi)) is actuated. This causes the engine and pitch controls tobe set in their proper position for the final glideslope portion of theapproach. In addition, this causes switch S₂ to close and S₁ to open.When these switches are activated, it causes h to be compared with H_(b)in decision blocks D₄ and D₅ (rather than the previous H_(a) value)Thus, the measured/calculated height will be compared to H_(b) forcontrolling the height during this phase of the approach.

The final decision block (D₆) compares the aircraft height h to thepre-determined height H_(ci) which occurs at point C in FIG. 2. This isthe point where the aircraft is just above touch down and all power isremoved and the aircraft nose is raised for touch down. Thus, a "no"response from decision 6 causes no action while a "yes" response causesall power to be removed, the nose to pitch up, and the approach isconcluded as the aircraft touches down on the surface.

The lateral control for the automatic landing system is performedsimilarly but the problem is more simple because it is only necessary toconstrain the aircraft to fly a desired heading from initial point A (ofFIG. 3) to point C where landing is accomplished. FIG. 3 gives thenecessary parameters to be measured and controlled by the yaw controlsystem.

As mentioned previously, there are several possible implementations forthe lateral (yaw axis) control of the air vehicle. One approach would beto command the air vehicle to fly directly toward the seeker so that (inthe yaw direction) it was operation about null. In this case, anyvariation of the seeker yaw angle from null (zero degrees) willrepresent an angular error and command a corresponding change inaircraft yaw position. This would have the advantage of simplifying theseeker gimbal structure (in the yaw direction) but it would require thatthe distance from the sensor/seeker to the air vehicle be extremely longat the beginning of the approach because the total landing sequence,including the landing roll, must be accomplished before the air vehiclereaches the seeker/sensor. Also, in the event that the air vehicletouches down at a higher than normal speed, it might skid into theseeker/sensor and cause great damage. Even more significant, the seekerpitch angles would have very small changes if the distance from theseeker to points B and C are great. Therefore, it is much more practical(and more simple to implement) if the seeker/sensor is placed slightlybeyond the desired touch down point and offset from the desired path bysome distance (such as 50 feet). Then, the position of the air vehiclewith respect to the desired lateral centerline can be geometricallycalculated by using the range r, the seeker yaw angle α, and the offsetdistance s.

The distance d_(m) (FIG. 3) is the same distance d_(m) utilized in FIG.2 for glideslope control. Thus when d_(m) is computed there it is alsoused in the lateral control system. The desired seeker yaw angle alpha(α) is computed as: ##EQU4##

Where S is chosen arbitrarily and d_(m) is computed in the glideslopeprocess (equation 3). Then the measured angle α_(m) can be directlycompared to the desired angle α_(d) and the following logic used forcontrol commands:

    if α.sub.m <α.sub.d ; right rudder (turn right)

    α.sub.m =α.sub.d ; no change

    α.sub.m >α.sub.d ; left rudder (turn left)

These commands would be very small corrections because the tracker iscontinuously calculating needed corrections and, therefore, errors wouldnot have an opportunity to become large. FIG. 9 illustrates a potentiallogical implementation of this process.

If it is desired to continue tracking the air vehicle for lateralcontrol after air vehicle touch down, the gimbal of the seeker is madeto allow 180 degree freedom in yaw so it can track around as the airvehicle passes and continue tracking as the air vehicle moves away.

A numerical example using FIGS. 3 and 9 will help to explain further theoperation of the directional control portion of the automatic landingsystem. The preselected value S (FIG. 7) is from the same source as thepreselected values of FIG. 8 (e.g. computer memory, terminal keyboard,etc.) and is chosen as:

    S=50 feet

Then (in FIG. 9), the desired yaw angle α_(d) is calculate in C9 usingd_(m) from FIG. 8 (glideslope control system) and the selected value ofS using equation 6. This 10 is compared with the measured yaw angleα_(m) from the seeker/sensor in decision block D₇. If α_(m) is less thanα_(d), a command is given to turn right. If the output of this decisionis "no", then a check is made (D₈) to see if α_(m) is greater thanα_(d). If this output is "yes", a left turn is commanded. If the outputis "no", then the aircraft is on the proper course and no action istaken.

Therefore, between the two sub-systems (glideslope and directionalcontrol) the aircraft can be automatically guided to a landing pointfrom an initial point in flight.

I claim:
 1. An automatic landing system for landing remotely pilotedflying vehicles in a predetermined flight path and at a predeterminedpoint, comprising;(a) an auto pilot carried by said flying vehicle formeasuring the parameters of attitude, airspeed, and heading andcomparing the same to predetermined parameters for the desired attitude,airspeed and heading, and for adjusting the path of said aircraftwhenever deviations from the desired parameters are detected; (b) radartransmitter and receiver means disposed on a stabilized gimbal formeasuring the angular deviation from a reference and the actual distancefrom said vehicle to said radar transmitter and receiver means on acontinuous basis; and (c) control means for receiving signals from saidradar, transmitter and receiver means indicating the pitch and yaw angleand the range from said radar transmitter and receiver means to the saidvehicle, computing both actual and desired altitude and heading angle ona continuous basis and comparing said desired and actual altitudes andheading on a continuous basis from the initiation of landing approachuntil said vehicle reaches a pre-determined touch-down point, and fortransmitting signals indicative of any deviation from said desiredparameters to said autopilot for correcting any such deviations fromsaid desired parameters whereby said remotely piloted flying vehicle islanded at a predetermined touchdown point and travels along apredetermined glideslope and heading during landing.
 2. An automaticlanding system as set forth in claim 1, wherein said radar transmitterand receiver means is offset from the landing path of said flyingvehicle.
 3. An automatic landing system as set forth in claim 1, whereinsaid radar transmitter and receiver means is supported upon a doublegimbal for measuring the yaw and the pitch angles that the radar mustmove through to track said vehicle.
 4. An automatic landing system asset forth in claim 1, wherein said radar transmitter and receiver ismounted on a portable base.
 5. An automatic landing system as set forthin claim 4, wherein said portable base comprises a pedestal.
 6. Anautomatic landing system as set forth in claim 4, wherein said basecomprises a tripod.